﻿Dnepr Rocket successfully Launches Cluster of 37 Satellites﻿

June 19, 2014

A Dnepr launch vehicle blasted off from Russia's Yasny Launch Base at 19:11 UTC on Thursday, embarking on a mission to deliver a record load of 37 active satellites of different shapes and sizes into orbit. Following liftoff out of a rocket silo at Site 370/13 at the Dombarovsky missile range, Dnepr started racing into the sky headed to an orbit at 630 Kilometers in altitude. The ascent mission took about 16 minutes and concluded with a precisely planned sequence to deploy the different payloads. Dnepr is a R-36M missile also known as SS-18 Satan that was stationed all across the Soviet Union starting in 1966 outfitted with multiple warheads and independent re-entry vehicles. After the end of the cold war and the fall of the Soviet Union, a portion of the R-36 fleet was modified to become Space Launch Vehicles. Dnepr is one of the cheapest launch vehicles that are currently flying offering a payload capability of up to 4,500 Kilograms to Low Earth Orbit. The three-stage rocket stands about 34.4 meters tall with a diameter of 3.0 meters and a mass of 211,000 Kilograms using only storable propellants - Unsymmetrical Dimethylhydrazine fuel and Nitrogen Tetroxide Oxidizer.

Countdown operations got underway about eight hours ahead of liftoff as the primary payloads were powered up for final checkouts, battery charging and reconfigurations, controlled by the spacecraft teams via data connections of the satellites.

*File Image* - Credit: ESA (Cryosat-2)

The launch on Thursday was threatened by a front of severe weather that passed through the area. As part of the standard pre-launch meeting of the Russian State Commission, weather was among the discussed topics along with the status of the launcher and ground systems. The storm front was expected to pass before launch time and no technical problems were present, allowing the formal GO for launch to be given.

The countdown for the Dnepr rocket started at T-3 hours - with its ICBM heritage, Dnepr does not require a lengthy countdown. The rocket was fueled in the days leading up to launch and final checkouts of its various systems were completed on Wednesday.

*File Image* - Credit: ESA (Cryosat-2)

*File Image* - Credit: ESA (Cryosat-2)

90 minutes ahead of launch, the silo door was opened and the launch area was evacuated after final hands-on work was completed at the silo facility. The onboard control system of the launcher was tested at T-1 hour before being deactivated again. At T-20 minutes, the control system was powered up again and final reconfigurations commenced leading to the final countdown phase starting three minutes before T-0. At T-3 minutes, the launcher switched to internal power and the launch system was enabled to trigger the liftoff sequence at the precise T-0 time. Ground station telemetry recorders were activated as well to prepare for liftoff.

The countdown hit zero at precisely 19:11:11 UTC and the silo's ejection system was activated. Dnepr uses a black powder mortar system that pushes a special tray at the base of the Dnepr rocket up to rapidly expel the rocket from its silo. As soon as Dnepr was out of the silo, the tray was ejected to the side by a small solid rocket motor in order to enable Dnepr to ignite its first stage engines.

A series of five O-Rings were pyrotechnically jettisoned at T+4 seconds and, at the same time, the RD-264 engine of the first stage ignited. RD-264 is a cluster of four RD-263 engines that share common turbopump equipment and provide a total thrust of 461,200 Kilograms – enough to lift the 211,000-Kilogram Dnepr launcher and its payloads hidden under the launcher’s fairing and gas shield.

Immediately after ignition, Dnepr started executing its pitch maneuver to begin heading south toward its polar target orbit, giving its payloads a smooth ride uphill.

Dnepr’s first stage measures 22.3 meters in length and 3 meters in diameter carrying 147,900 Kilograms of propellants used by the main engines which operate at a high chamber pressure of 220 bar to generate 4,523 Kilonewtons of vacuum thrust. Control during first stage flight was provided by individually gimbaling the engine nozzles.The first stage shut down and separated from the second stage at approximately T+1 minute and 38 seconds. Upon stage separation, the second stage ignited its RD-256 main engine and RD-257 vernier achieving a total thrust of 77,000 Kilograms. The second stage is 5.7 meters long and carries 36,740 Kilograms of propellants. Control during second stage flight is provided by the vernier engine that features four individually gimbaled nozzles.

While the second stage was burning, the launcher jettisoned the upper portion of its payload fairing that had protected the satellites during the initial portion of the ascent when aerodynamic forces would have damaged the payloads. For Dnepr, the fairing is only one of two protective devices, the second being the gas dynamic shield that protects the delicate satellites on the upper deck from engine plumes of the third stage.The second stage performed a burn of approximately two minutes and 48 seconds before separating from the third stage of the launcher.

After staging, the third stage was in a short coast phase to climb up near the apogee of the trajectory so that its burn could act as a circularization maneuver and raise the perigee. The mission was targeting a circular sun-synchronous orbit at an altitude of 630 Kilometers and an inclination of 98°.

The third stage of the Dnepr launcher uses four RD-869 engines that are firing forward - towards the nose of the launch vehicle. This unusual propulsion scheme is a relic of the original R-36 concept that required the upper stage to perform multiple warhead deployments in a short time frame.

The engines swung out after stage separation and ignited when the coast was complete - starting the burn in a high thrust mode and immediately beginning a 180-degree flip of the vehicle to start providing positive thrust to the stack - pulling the stack into orbit instead of pushing the vehicle as most rockets do.

The four RD-869 engines deliver 2,070 Kilograms of thrust at nominal throttle and 846kgf in throttled mode. Overall, Dnepr’s third stage is 1 meter long and uses a propellant load of 1,910 Kilograms. Control is provided by gimbaling the engines that continue to fire throughout the primary mission - including the payload separation sequence.

When the third stage approached its target orbit, the Gas Dynamic Shield was jettisoned to expose the payloads for the precisely planned separation sequence controlled by a central sequencer that is part of the space head module and interfaces with the various payload adapters and deployment mechanisms.

For this Cluster Mission, the SHM-2 design was used that includes two decks and a bridge to facilitate all the different payload adapters and CubeSat Deployers.

BugSat-1, an experimental Earth observation satellite, was the first satellite to be separated at T+15 minutes and 50 seconds. With the third stage still firing, the space head module pulled away from the deployed satellite to create the proper separation distance for the next separation which was TableSat-Aurora at T+15:52, a Russian technology demonstration satellite.

The Deimos-2 Satellite was third to be released from its payload adapter at T+15 minutes 54 seconds. Deimos-2 hosts the High Resolution Advanced Imaging System for operation by Elecnor Deimos to acquire high-resolution images of Earth for a variety of purposes.

Image: Kosmotras

Image: Kosmotras

Deimos-2 was followed a short time later by the carefully sequenced deployment of 21 CubeSats from five QuadPack deployers which included the release of 11 Flock-1c satellites to become part of the Flock-1 Earth observation constellation (detailed payload list below).

Once the Deck-A payloads were released, the payload platform was jettisoned to expose the Deck-B payloads for deployment as the third stage continued firing to establish spacing between the objects. 13 satellites of different sizes were then released from that platform in close succession including the KazEOSat-2 Earth Observation satellite (T+15:58), the Japanese Hodoyoshi twins that will demonstrate optical imaging systems (T+16:00 & T+ 16:01), the SaudiSat-4 Gravitational Reference Sensor demonstrator satellite at T+16:02.5, the AprizeSat-9 and 10 pair of communication sats (T+16:05.5 & T+16:07) and the BRITE-CA 1&2 pair to study bright stars (T+16:10 & T+16:12). Also part of the second batch of satellites to be deployed was UniSat-6 (at T+16:04) which in itself contains four CubeSats for deployment 24 hours after launch: the 3U Lemur-1 and TigriSat, the 2U ANTELSAT and the 1U AeroCube6.

The entire satellite deployment sequence took less than 30 seconds and was completed at about T+16 minutes. Following the successful payload deployment, the RD-869 engines continued to fire to move the stage away from the satellites in order to place it into a different orbit to avoid re-contact with the multitude of payloads and finish a successful Dnepr launch.

Payload Overview - 37 Satellites

﻿KazEOSat-2﻿

Photo: SSTL

Image: SSTL

KazEOSat-2 is a medium-resolution optical Earth Imaging Satellite built by Astrium Defence and Space and Surrey Satellite Technology for operation by the Kazakh Space Agency to provide data to the government of Kazakhstan to support decision-making and policy in critical area such as resource monitoring, resource management, land-use mapping and environmental monitoring information.

Two KazEOSat spacecraft are launched in 2014, first was KazEOSat-1 also known as DZZ-HR or HRES (High Resolution Earth Observation Satellite) that was delivered to orbit by the European Vega launcher on April 30, 2014. While KazEOSat-1 provides high-resolution imagery, it will be complemented by the lower resolution images of KazEOSat-2 that cover a wider swath and thus reduce revisit times. The KazEOSat-2 satellite is also known as DZZ-MR or MRES (Medium Resolution Earth Observation Satellite).

The KazEOSat-2 spacecraft is based on the SSTL-150 small satellite platform that can host a range of optical imaging payloads. Overall, KazEOSat-2 weighs 185 Kilograms and is 70 by 80 by 90 centimeters in size.

The imaging payload uses a Three Mirror Anastigmatic design that, provides a larger field of view which is a requirement for the KazEOSat-2 satellite to achieve a wide ground swath. The telescope uses aluminum structures installed on a stable optical payload bench. The KEIS telescope aperture diameter is 145 millimeters allowing light to pass onto the primary, secondary and tertiary mirrors that focus the light into the bandpass filters and detector unit. The optical system uses a folded design without optical obstructions to enable it to take a small volume and keep the aperture small. Overall, the telescope has an effective focal length of 633 millimeters.

﻿Deimos-2﻿

Deimos-2 is a Spanish Earth Observation Satellite as a follow-on to Deimos-1 that launched in 2009. The satellite was designed and developed by Satrec Initiative, an Elecnor company, for operator Deimos to acquire high-resolution images of Earth for a variety of purposes.

Deimos-2 uses a multispectral pushbroom imager to capture images on a 12-Kilometer swath achieving a resolution of one meter for panchromatic images and four meters in the visible and near infrared spectral range.

The satellite is based on the SI-300 spacecraft bus that is a small-class platform capable of hosting Earth observation and science payloads. Features of the platform are hall-effect thrusters for orbit control and maintenance, agile and accurate pointing for precise imaging and a dual-redundant systems architecture. Deimos-2 is 1.5 by 1.95 meters in size with a mass of ~310 Kilograms.

The High Resolution Advanced Imaging System was designed and developed by SI (Satrec Initiative) and Elecnor Deimos and has successfully flown on previous missions. The payload consists of three main elements – the electro-optical subsystem, the solid state recorder and the Image Transmission Unit.

The payload features a Korsch-type telescope with a five-mirror design with a field of view of 1.2 degrees. The main mirror of the satellite is 41.5 centimeters in diameter. Three folding mirrors are used to create a total focal length of 5.7 meters. The imaging payload covers a panchromatic channel (450-900nm) and four multispectral channels – Blue (420-510nm), Green (510-580nm), Red (600-720nm) and Near Infrared (760-890nm). Multispectral resolution is 4 meters and the sharpened-PAN mode achieves ground resolutions of up to 0.75 meters.>>>Detailed Satellite Overview

Image: SI, Elecnor Deimos

Photo: SI, Elecnor Deimos

﻿11 Flock 1c Satellites﻿

Photo: NASA

Image: Planet Labs

Image: Planet Labs

Flock 1 is a satellite constellation of CubeSats dedicated to Earth Observations using a fleet of small satellites to generate high-resolution images of Earth achieving resolutions of three to five meters. The operational constellation began deployment in 2014 and uses a combination of shorter and longer lived orbits being launched from the International Space Station and different orbital launch vehicles.

The satellites are designed, developed, manufactured and operated by Planet Labs based in San Francisco that markets the Earth Observation data products to a range of customers for a variety of applications. The Flock 1 spacecraft are based on the three-unit CubeSat specification having a launch mass of about 5 Kilograms and being 100mm × 100mm × 340mm in size featuring body mounted solar panels and two deployable solar arrays with three panels each using triangular advanced solar cells.

Attitude data is provided by three-axis magnetometers to accomplish three-axis stabilization via a reaction wheel system and magnetic torquers for momentum management. Fine pointing data is provided by a Star Camera. Flock 1 satellites use a single-board computer to control all spacecraft and payload functions with a watchdog board able to reboot the flight computer in the event of errors or radiation related upsets.

The satellites use an X-Band system for the downlink of acquired images and systems telemetry at data rates of up to 120Mbit/s. Primary command uplink is done via S-Band, although a low-speed Telemetry and Command System operating in the UHF band is also available and in use for early commissioning operations and as a backup.

The main payload of each satellite is an optical telescope of unknown specifications to acquire high-resolution images of Earth. The telescope has an aperture diameter of 90mm and is protected by an aperture cover that is deployed via springs. The optical axis is down the central axis of the satellite to achieve a maximum focal length. >>>Flock 1 Mission Page

﻿SaudiSat-4﻿

SaudiSat-4 is a Saudi-Arabian technology demonstration satellite that was developed by the King Abdulaziz City for Science & Technology with input from NASA Ames to conduct a UV LED Experiment in space that will demonstrate UV LED technology and also study AC charge management in LEO satellites. SaudiSat-4 weighs approximately 112 Kilograms based on a modified version of the SaudiSat-3 satellite bus with a mass of 83 Kilograms. The spacecraft uses four fixed solar panels to generate 61 to 86 Watts of electrical power that is distributed to the various satellite subsystems including the payload that requires 45 Watts of power during operation. Batteries are used to store up to 450 Watt-hours of electrical power. The satellite is three-axis stabilized using two star trackers, magnetometers and sun sensors to acquire orientation data for precise pointing using a reaction wheel assembly. SaudiSat-4 has no propulsion system for orbit control. The spacecraft uses an integrated avionics unit that houses the onboard computer and other avionics that control all satellite functions and execute commands sent from the ground, actuating all subsystems and commanding payload operations. SaudiSat-4 has a mass memory of 25Gbytes that can accept data are rates of up to 25Mbps. The satellite uses an X-Band communication system achieving a downlink data rate of up to 358 Mbit/s. The heart of the satellite is its payload – a Modular Gravitational Reference Sensor and a UV LED Charge Management System for demonstration in orbit. The system consists of a proof mass – a 700-gram sphere that is coated with gold, four bias plates, 16 UV LEDS, a contact probe and gold-coated Ultem tubes for shielding. The Gravitational Reference Sensor using a free-floating test mass on drag-free spacecraft requires the control of charge build-up on the proof mass. The GRS is an accurate accelerometer that is limited by the charge build-up on the free floating test mass. A GRS is used to allow a spacecraft to fly in an orbit that is solely influenced by gravity and not affected by solar pressure and atmospheric drag. This type of drag-free flight scheme is utilized in gravity sensing spacecraft.

Image: KACST

Image: NASA/Stanford University

Image: NASA/Stanford University

Image: NASA/Stanford University

Image: NASA/Stanford University

Image: NASA/Stanford University

The UV LED System is designed to remove the charge build-up on the test mass by exposing it to UV illumination which creates a cloud of electrons as a result of photoemission leading to the discharge of the sphere. The discharge rate is controlled with electrodes that are biased in voltage. Using the LEDs and bias in phase (0°), photoelectrons are transported from the test mass to the electrodes creating a positive charge transfer. When the LEDs and voltage bias are operated out of phase (180°), photoelectrons are transported from the electrodes to the test mass as part of a negative charge transfer. The LEDs have a wavelength maximum at 255 nanometers.

The displacement of the test mass is measured by an Differential Optical Shadow Sensor that uses eight LEDs and photodiodes on the opposite side of the cage of the test mass are used to determine the displacement of the test mass as a function of light intensity measured by the photodiodes. The DOSS system achieves a 10-nanometer resolution at a frequency of up to 1Hz.

The caging mechanism has been designed for a 200-Newton holding force to restrain the sphere during launch.

SaudiSat-4 will qualify the UV LEDs for spacecraft and demonstrate the AC charge management system. The success of the GRS will increase the capability of gravitational reference sensors for future missions that are currently under development such as the Laser Interferometer Space Antenna (LISA). Future nano and microsatellite missions will carry Modular Gravitational Reference Sensors using different variations such as grating angular sensing and diffractive optics as well as micro and nanothrusters to eventually develop the MRGS for scientific missions.

﻿Hodoyoshi 3 & 4﻿

The Hodoyoshi 3 and 4 satellites are two Earth Observation satellites developed and operated by the University of Tokyo to demonstrate a series of satellite systems with particular focus on the optical imagers that are being tested on the spacecraft. Both satellites use a similar structure, but include different subsystems for technical demonstrations.

Image: University of Tokyo

Hodoyoshi 3 is 0.5 by 0.5 by 0.65 meters in dimensions weighing 60 Kilograms. The spacecraft is equipped with body mounted solar cells and two deployable solar panels to deliver a maximum power of 100 Watts that is distributed using a 28V and a 5V power bus. Power is stored in a 5.8Ah Li-Ion battery and the expected average power consumption of all satellite subsystems is 50 Watts. The satellite is three-axis stabilized using star trackers, an inertial measurement unit and a reaction wheel assembly to provide very accurate Earth pointing capabilities. Telemetry downlink and command uplink is accomplished via an S-Band system with an uplink rate of 4kbps and a downlink rate of 4 to 64kbps. Mission data is downlinked via an X-Band system achieving data rates of up to 10 Mbps. The Hodoyoshi-3 satellite is outfitted with a Hydrogen Peroxide propulsion system for orbital maneuvers and orbit maintenance.

Hodoyoshi-3 is equipped with a Low-Resolution Camera that can achieve ground resolutions of 200 meters along a ground swath of 491.5 Kilometers and a medium resolution imager that can reach a ground resolution of 40 meters with a swath width of 82 Kilometers.

The Hodoyoshi-4 satellite is 0.5 by 0.6 by 0.7 meters in size with a mass of 66 Kilograms. It also uses body-mounted solar panels and two deployable panels to generate 100 Watts of power. Satellite stabilization is also accomplished with star tracker data and reaction wheels. Communications are done in S-Band for TT&C like on Hodoyoshi-3, but the Hodoyoshi-4 satellite uses an improved X-Band system that will demonstrate data rates of up to 100Mbps.

The Hodoyoshi 4 satellite is equipped with an Xenon-fueled ion-thruster for propulsion achieving thrust of 260 micronewtons and a specific impulse of 1170 seconds. The system weighs 6 Kilograms and measures 39 by 28 by 16 centimeters and includes a Xenon tank with a capacity of about one Kilogram. The propulsion system can deliver a total delta-v of 240 m/s requiring a peak power of 30 Watts.

The Hodoyoshi-4 spacecraft is also demonstrating an Earth observation payload. This high-resolution optical imager uses a pushbroom sensor to acquire Earth imagery at a ground resolution of 5 meters. The system weighs 3.5 Kilograms and has an aperture diameter of 15 centimeters and a focal length of 10 centimeters. The imaging payload covers four spectral bands – Blue (450-520nm), Green (520-600nm), Red (630-690nm) and Near-Infrared (730-900nm). The observable ground swath is 30 Kilometers wide.

Both Hodoyoshi satellites carry a secondary Store-and-Forward Payload for water level monitoring sensors deployed in many places around the globe. The store and forward payload operates at a UHF frequency of 400MHz and uses sampling frequencies of 10kHz or 40Hz. The payload has an onboard data storage capacity of up to 16Gbits to store data packets that are received in data transmissions of one second transferring 270bits and ten seconds transferring 2970bits of data. Using the two satellites, the monitoring sensors have passes every 11 to 13 hours, every 24 hours in the worst case. Data latency from the point of acquisition to downlink to a ground station is under six hours.

﻿TableSat-Aurora﻿

TableSat-Aurora, also known at TableSat-Aurora-2U-EO is a 25-Kilogram satellite designed and developed by SPUTNIX, a Russian start-up company specialized in the design of small satellites up to 50 Kilograms for Earth observation, communication and technical demonstrator payloads.

The spacecraft uses a hexagonal satellite bus that features six deployable fixed solar panels for power generation. Power is stored in LiFePO4 batteries and distributed by a dedicated avionics unit. The satellite will test a PCB-based GaAs solar panel. TableSat-Aurora is three-axis stabilized using star trackers, sun sensors, a magnetometer and an inertial measurement unit for the acquisition of precise orientation data. Fine pointing is achieved through the use of reaction wheels and magnetic torquers for momentum dumps. The SPUTNIX sun sensor has a field of view of +/-60 degrees for accurate determination of the sun vector. It is 19 by 46 by 38mm in size and weighs 100 grams requiring 0.05W of power. The SX-SR-Mini-01 Star Tracker obtains attitude information with an accuracy of under 1” with an update rate of up to 10 Hz to be capable to track stars up to a spacecraft angular velocity of 3deg/s.

The spacecraft is equipped with a GPS receiver, a UHF terminal for command uplink and telemetry downlink and an X-Band transmitter for data downlink.

The X-Band terminal uses switched array technology providing a wide scan sector of 128 degrees to allow the satellite to maintain its nominal attitude during communication sessions without the need of maneuvers to track a ground station. The system operates at a frequency of 7.95 to 8.6 GHz.

TabletSat-Aurora uses a Pug-n-Play architecture based on the SpaceWire bus for all of its onboard systems to simplify systems integration and the overall architecture of the spacecraft to create a flexible satellite design.

The satellite carries an Earth imaging payload equipped with a panchromatic photo- and video- camera that covers a spectral range of 430 to 950 nanometers. The payload covers a swath of 47 Kilometers and achieves a ground resolution of 15 meters.

Image: SPUTNIX

Image: SPUTNIX

﻿BugSat-1﻿

Photo: ISIS Launch Services

BugSat-1, also known as Tita, is an Argentinean technology demonstration satellite developed and operated by Satellogic S.A. The 22-Kilogram satellite hosts a medium-resolution optical imaging payload that will be tested as part of its mission. The spacecraft also includes amateur radio equipment for outreach activities after the demonstration mission is complete. Overall, the spacecraft is 27.5 by 50 by 50 centimeters in size and includes body-mounted solar panels and a three-axis stabilization system for pointing to Earth required for the acquisition of ground images. The spacecraft uses a GPS receiver for orbit determination, a UHF communications system and a C-Band radio for telemetry downlink in the UHF band. The spacecraft tests a medium-resolution optical imaging system capable of acquiring images in multiple spectral bands. The design of the imaging payload will be verified in orbit to provide information for the design of the follow-on satellite, BugSat-2 and eventually an operational constellation of Earth imaging spacecraft. Once the demonstration mission of the optical imager is complete, BugSat-1 will transition to an outreach mode in which amateur radio operators can receive payload data and use an onboard Digipeater to relay radio messages.

﻿AprizeSat 9 & 10﻿

AprizeSat 9 and 10 are part of a constellation of Low Earth Orbit satellites that provides global communication services for data transmission and fixed and mobile asset tracking and monitoring including machine to machine communications. The satellites are cubical in shape with a side length of 25 centimeters using the typical microsatellite design consisting of different trays connected by a cable harness.

The side panels are sheet metal and host the body-mounted solar panels – the four side panels generate 15 Watts of power while the top and bottom panels generate 7.7W each. The satellites uses Ultra Triple Junction Gallium-Arsenide cells with high efficiency feeding Nickel-Cadmium batteries. Peak Power Point Tracking is employed and the satellite uses regulated 3.3V, 5.5V and 12V supplies in addition to an unregulated 8V bus. The satellites use permanent magnets to keep aligned to the magnetic field of the Earth and a series of alternating strips of reflective and non-reflective tape induce a roll of about 0.25rpm to distribute sunlight on the exterior of the satellite. The spin rate is limited by magnetic hysteresis rods.

The 12-Kilogram satellites carry ten radio receivers, two power-agile transmitters and 12MB of solid sate data storage. The radios operate in the VHF/UHF frequency band to collect data from more than 100,000 user terminals each day with data rates of 1.2 to 38.8kbit/s. An S-Band terminal supports data rates of several Mbit/s for up and downlink for users with large payload data requirements.

Photo: SpaceQuest

These are the 10th and 11th satellites to be launched since 2002. The expected satellite lifetime is 10 to 12 years and the planned constellation will consist of a total of 64 satellites. AprizeSat-10 carries an Automatic Identification System Terminal. The Automatic Identification System is used by sea vessels that send and receive VHF messages containing identification, position, course and speed information to allow the monitoring of vessel movements and collision avoidance as well as alerting in the event of sudden speed changes. These signals can be transmitted from ship-to-ship and ship-to-shore to allow the monitoring of a local area, but deploying space-based AIS terminals allows a broad coverage and data relay to ground stations for monitoring of large sea areas.

The system consists of the Ground Primary Payload System GPPL which is comprised of small transmitters that are mounted on birds to allow tracking from the Primary Payload subsystem on the DTUSat-2 spacecraft that is capable of receiving and tracking the 868MHz signals of the GPPL transmitters. The GPPL to PPL data link uses BPSK modulation for the transmission of small data packets that allow the identification of the bird and its position. The GPPL system uses a dipole antenna on the +Z face of the satellite and the digitized signal is provided to the onboard computer for storage and downlink to ground stations.

The GPPL ground based transmitters implement the latest miniaturization technologies to create an ultra small unit. GPPL uses a single 30mm² ultra triple junction solar cell with 28% efficiency to charge a battery to sufficient power levels for transmission during the migration payload. The GPPL use a 16-bit processor to propagate the orbit of the DTUSat-2 satellite to determine the time of ground passes for the exchange of data. This calculation is based on time and position solutions of a mini GPS receiver. When the satellite is in view, GPPL will activate the BPSK transmitter and start transmitting data at 30bit/s containing the last five position fixes that were stored onboard the sensor.

DTUSat-2 is outfitted with a Pico Camera that uses a CCD detector operating in the visible wavelength and using a Bayer filter to obtain color images. The camera has a resolution of 640 by 480 pixels with an 8-bit gray scale to be used to monitor the gravity gradient boom deployment and to obtain images of stars to check the performance of the attitude determination system.

Another payload tested on DTUSat-2 is a second generation sun sensor using a Micro-Opto-Electro-Mechanical-System delivering an angular accuracy of under one degree.

﻿UniSat-6﻿

The UniSat-6 spacecraft was developed by GAUSS Srl, Rome using the UniSat satellite platform that is 50 by 50 by 50 centimeters in size using reinforced aluminum and carbon honeycomb panels. The spacecraft weighs 26 Kilograms. Body-mounted solar panels provide 30 Watts of power to the satellite, attitude control is accomplished through the use if permanent magnets and a UHF communication system is used for command uplink and data downlink.

UniSat-6 includes two Pico-Orbital Deployers and one PEPPOD (Planted Elementary Platform for Picosatellite Orbital Deployment) for the release of four CubeSats from the spacecraft. These four satellites are the 3U Lemur-1 and TigriSat, the 2U ANTELSAT and the 1U AeroCube6. The deployment of the satellites is triggered 25 hours and 34 minutes after launch using a completely autonomous system with a dedicated battery. A backup system is also included that can deploy the satellites using ground commanding. A camera will acquire imagery of the separation of the CubeSats.

Photo: GAUSS

﻿BRITE-CA 1 & 2

Photo: University of Toronto

Image: University of Toronto

The BRITE-CA 1 and 2 satellites are also known under their full name of BRIght-star Target Explorer or CanX-3 (Canadian Advanced Nanospace eXperiments). These two 7-Kilogram spacecraft host a wide-field telescope payload that will be used to conduct differential photometry measurements of bright stars.The BRITE-CA satellites were developed at the University of Toronto to follow the UniBRITE, BRITE-Austria and BRITE-PL satellites that were funded before the two Canadian contributions received funding.

The BRITE-CA satellites use the Generic Nanosatellite Bus designed around a 20 by 20 by 20 centimeter cube. Power is provided by four to ten triple-junction GaAs solar cells installed on each of the external panels delivering up to ten Watts of power. Power is stored in Li-Ion batteries with a capacity of 5.3 Ah and the power conditioning unit provides a 4-Volt unregulated power bus. Attitude Determination is accomplished by a three-axis magnetometer, six sun sensors for fine and sun attitude determination and a star tracker for precise attitude determination. The Miniature Star Tracker provides three-axis attitude solutions at a control cycle at 0.5 Hz and an accuracy of 10arcsec. Attitude actuation is provided by three reaction wheels with a total mass of 185grams and a volume of 5 by 5 by 4 centimeters. The wheels have a momentum capacity of 30mNms and deliver a maximum torque of 2mNm. Momentum dumps are supported by three magnetotorquers.

Data handling and satellite control is provided by an ARM7 housekeeping computer that handles standard telemetry and communications while a second computer supports all attitude determination and control functions. A third computer board is in charge of the operation of the science payload and handles its data. Each processor board uses the ARM7/TDMI processor with a code memory of 256kB and 2MB of hardware SRAM memory used to store program variables and data. A 256MB flash memory is used for long-term data storage.

The communications system of the satellite uses an S-Band system operating at 2234.4 MHz for data downlink reaching data rates of 32 to 256kbit/s. Uplink is accomplished at 437MHz in UHF with a data rate of 4kbit/s. A 0.1W VHF 145MHz beacon can be used for spacecraft tracking. BRITE-CA will downlink up to 8MB of payload data per day.

Image: University Graz

The science payload of BRITE-CA includes a photometer consisting of an optical head, an electronics tray and the baffle. The baffle includes an aperture stop and filters while the optical head houses five lenses. BRITE-CA examines the brightest stars in the sky for variability using precise differential photometry over time scales of days and months to answer questions on the life cycles of bright stars.

Using its payload, BRITE is capable of studying all stars brighter than magnitude +3.5. To track the periodic, semi-periodic and irregular variations of bright stars that occur on time scales of minutes to months, BRITE will perform continuous precise photometric time-series measurements by making measurements of a target star field of at least 15 minutes per orbit over a period of several months. Changes in star intensity occur due to changes in density, magnetic field, surface temperature and internal seismic phenomena.

To capture data on a wide field of view, the BRITE photometer has an FOV of 24 by 19°.

The photometer payload uses a five-lens design with an aperture of three centimeters and a CCD detector. The optical cell uses spacers to hold the lenses in their positions with a total focal length of seven centimeters. Overall, the photometer weighs 900 grams and requires 3W of power,

The detector is a KAI-11002 CCD – a 11 megapixel interline, buried channel CCD with 4008 by 2672 pixels that are 9 by 9 micrometers in size creating an image size of 37.25 by 25.70 mm. An electronic shutter is implemented in the system as well as anti-blooming protection. The effective wavelength range of the instrument is limited by the sensitivity of the detector in the red spectral range and in the blue by the transmission properties of the lenses. The optical filters were designed so that stars of 10,000K cause the same detector output. The blue filter installed on one BRITE-CA spacecraft covers a spectral range of 390 to 460nm while the red filter of the other satellite covers 550 to 700nm with a transmission rate of 95%.

The photometer is capable of performing differential photometry measurements with an error of less than 0.1% over a 15-minute observation using exposure times of 0.1 to 100 seconds. The CCD read-out is converted from analog to digital using a 14-bit scheme to convert the analog pixel values. A 32MB instrument memory can temporarily hold a full picture until it can be transmitted to the spacecraft.

Image: University Graz

Photo: University of Toronto

﻿Perseus M1 & M2﻿

Photo: Dauria Aerospace

Perseus M1 and M2 are two identical 6-Unit CubeSats developed and operated by Dauria Aerospace. The spacecraft carry and AIS payload. The Automatic Identification System is used by sea vessels that send and receive VHF messages containing identification, position, course and speed information to allow the monitoring of vessel movements and collision avoidance as well as alerting in the event of sudden speed changes. These signals can be transmitted from ship-to-ship and ship-to-shore to allow the monitoring of a local area, but deploying space-based AIS terminals allows a broad coverage and data relay to ground stations for monitoring of large sea areas. However, due to the large footprint of satellites, overlapping and signal collisions become a problem, especially for frequented traffic routes.

QB50P1 & QB50P2

QB50P1 and QB50P2 are two precursor satellites to the QB50 project that will launch a network of 50 satellites by a team of 15 universities and institutions around the world. The goal of QB50 is to demonstrate that a wealth of CubeSats can be launched by a low-cost launch vehicle to establish a constellation of scientific spacecraft in Low Earth Orbit enabling thermospheric science. The project will demonstrate the harmonization and coordination between many participating institutions and also establish an unprecedented scientific satellite constellation for in-situ measurements within the thermosphere, a zone that is hardly accessed by any other spacecraft for dedicated measurements. Three different types of scientific payloads will eventually be part of the QB50 spacecraft consisting of an Ion-Neutral Mass Spectrometer (INMS), the Flux-Φ-Probe Experiment and multi-Needle Langmuir Probe (m-NLP). These payloads will be combined with thermistors and thermocouples. The two precursor satellites will conduct demonstration mission as part of the risk reduction process for the whole QB50 project. The deployment System consists of a 3U QuadPack that has a total of 12-Units of launch capacity in 4 individually controlled launch-tubes and the associated deployment sequencing electronics. Using the deployment system on the precursor flight provides flight heritage as it goes through a ground-processing flow and an actual launch operation. The satellite platform and instruments will also undergo risk reduction as the spacecraft are put through six months of rigorous testing in orbit that can expose any shortcomings in systems performance before the operational constellation is launched. The QB50P1 satellite is a 2U CubeSat that is hosting the INMS payload from MSSL, the attitude determination and control system from SSC, the Thermocouple experiment from VKI and an AMSAT-NL transponder. The INMS (Ion-Neutral Mass Spectrometer) measures dominant species in the thermosphere such as atomic oxygen, molecular oxygen, nitrogen and nitrous oxide. The instrument also provides density and temperature measurements. INMS includes an ion filter, an ionizer unit and an electrostatic analyzer with an energy range of 0.1 to 28eV. The instrument achieves an elevation resolution of 5° and an azimuth resolution of 5° at a sample time of 4ms and energy resolution of under 3%. The QB50P2 satellite is also a 2U CubeSat with a launch mass of 2 Kilograms. It is identical to the QB50P1 satellite except for using the FIPEX (Flux-(Phi)-Probe-Experiment) instrument instead of the INMS.

Photo: ISIS

Image: QB50 Project

Image: QB50 Project

The FIPEX microsensor uses a gold cathode for non-dissociative reactions to detect atomic oxygen and a platinum anode for dissociative adsorption to detect atomic and molecular oxygen. The sensor system is 36 by 30 by 12 mm in dimensions and offers a field of view of about 180 degrees. The electronics unit controls the instrument and handles data coming from the two sensors to generate time resolved data for the measurement of oxygen and correlation to atmospheric models.

In addition to the standard instruments, the two satellites also host an amateur radio terminal operating in the VHF/UHF bands. After six months of rigorous testing, the satellites will be handed over to the amateur radio community for continued operation.

﻿PACE﻿

Image: NCKU

Image: NCKU

PACE - Platform for Attitude Control Experiments – is a project of National Cheng Kung University, Taiwan that was initiated in 2003. The 2U CubeSat is 112 by 112 by 243 millimeters in size and weighs less than 2 Kilograms hosting an experimental attitude determination and control system for testing in orbit.

The spacecraft consists of an aluminum framework and panels. 20 body-mounted triple-junction GaAs solar cells are used to generate two watts of power that is stored in Li-Ion batteries with an capacity of 3.6Ah. PACE uses a 3.3 and 5V power bus. Communications are accomplished via UHF using the 437MHz frequency. The satellite is controlled by an 8051 microprocessor that uses 16MB of flash memory for data storage.

The Attitude Determination and Control System of the PACE satellite consists of four digital sun sensors, a three-axis magnetometer and three gyroscopes for precise attitude determination and a momentum wheel assembly and magnetic coils for three-axis attitude actuation. The wheels spin at 2000rpm with a momentum capacity of 0.002Nms, the Digital Sun Sensors have a field of view of +/-60 degrees and provide an accuracy of 1 degree.

Throughout its mission, the satellite will downlink attitude information to evaluate the performance of the attitude control system. Using many newly developed components, PACE will also demonstrate these various systems in the space environment for potential use on future missions. Over the course of the mission, PACE will run different attitude control laws that are uplinked during the mission and executed sequentially. After attitude testing is complete, the satellite enters an automatic mode to study the behavior of its systems in the space environment over an extended period of time.

﻿POPSAT-HIP-1﻿

The POPSAT-HIP1 satellite was built by Microspace Rapid Pte Ltd., Singapore and uses the 3U form factor. It demonstrates a high-resolution optical imaging payload and an attitude control and propulsion system on a CubeSat class payload. The satellite operates in the amateur UHF band at 437 MHz.

Photo:
ISIS Launch Services

﻿Lemur 1﻿

Lemur 1 is a 3U CubeSat built and operated by NanoSatisfi to perform a technology demonstration mission of various satellite systems for use on future spacecraft. The satellite carries two Earth-observation payloads featuring electro-optical detectors in the visible spectral range reaching ground resolutions of five meters. The secondary payload is a low-resolution infrared imager with a ground resolution of one Kilometer. Lemur-1 is deployed from the UniSat-6 spacecraft after separation from the launch vehicle.

﻿NanoSatC-Br 1﻿

Image: INPE

Image: INPE

NanoSatC-Br 1 is a Brazilian CubeSat project managed by NPE (Instituto de Pesquisas Espaciais) South Regional Center. The small 1.3kg 1U CubeSat is equipped with a scientific payload but is also set fro a technology demonstration mission. The satellite uses the standard CubeSat features – body-mounted solar panels, Li-Ion batteries, a basic attitude determination and control system and processor cards to control all spacecraft functions. Communications are accomplished in the UHF/VHF bands for telemetry and command exchange while an S-Band systems is used for payload data downlink to ground stations at INPE and São José dos Campos.

The satellite is equipped with a Fluxgate Magnetometer Payload to measure the intensity of Earth's magnetic field over the South Atlantic Anomaly and the Equatorial Electrojet. The magnetometer records several data points per second collecting three-axis magnetometer data at an intensity resolution of 15 Nanotesla. The hall sensor can support a magnetic field range of 63mT. Overall, the sensor is 8.9 by 8.9 by 3.2 millimeters in size sampling at 5kHz. The second payload is a particle precipitation dosimeter to measure the radiation environment in Low Earth Orbit as a function of magnetic field strength.

The South Atlantic Anomaly is of particular interest because of a decreased shielding of Earth’s magnetic field causing high proton fluxes which can lead to radiation related upsets in satellites and is also of importance to air traffic. The Equatorial Electrojet is a flow in Earth’s ionosphere to the east near the equator. The EEJ causes daily variations in magnetic field strength and has been the subject of several space missions.

In addition to the science payloads, the satellite is equipped with a Field Programmable Gate Array and Integrated Circuitboard for testing in the radiation environment found in Low Earth Orbit. These components will be monitored over time to identify any deficiencies that may be caused by radiation.

﻿Aerocube 6﻿

Aerocube 6 is part of the small satellite program of The Aerospace Corporation, El Segundo, California. The AeroCube program deploys small satellites for technical demonstrations. Aerocube 6 is a 0.5U CubeSat that incorporates a host of technology upgrades and a suite of micro dosimeters to measure the radiation environment in Low Earth Orbit. The satellite is spin-stabilized and constantly sun-pointing. It includes a inter-satellite cross link, an integrated flight computer, a GPS receiver and a UHF transceiver unit. Aerocube 6 is deployed from the UniSat-6 satellite.

﻿PolyITAN-1﻿

PolyITAN-1 is a 1U CubeSat designed and developed at the National Technical University of Ukraine to perform a number of test and demonstration objectives in orbit. The satellite uses CFRP (Carbon Fiber Reinforced Plastic) panels and structures and the common CubeSat features such as body-mounted solar cells, a basic attitude determination and control system and a VHF/UHF and S-Band communication systems. The attitude determination system uses a three-axis magnetometer, a three-axis gyro and five sun sensors and attitude control is accomplished using three electromagnets to align the three satellite axes in Earth’s magnetic field. The satellite uses a navigation system compatible with Glonass and GPS satellites for precise orbit determination.

PolyITAN-1 allows university students to build, test and launch a small satellite to gather valuable experience. The mission’s technical demonstrations include the sun sensors, the attitude determination software and the miniaturized Glonass/GPS receiver. PolyITAN-1 also tests a temperature probe, a dust sensor and an Ultraviolet sensor to track changes in performance over the course of the mission.

Photo: KPI

Furthermore, the satellite system implements mini heat pipes that are being evaluated for future thermal control systems of small spacecraft. The mission also includes the development and improvement of ground systems architecture in preparation for future small satellite missions.

﻿ANTELSAT﻿

Photo: FING

Antelsat is an Uruguayan CubeSat using the 2U form factor. The satellite has been designed and built at Facultad de Ingeniería de la Universidad de la República (FING) to allow students to develop skills in radio and aerospace engineering. The spacecraft is equipped with cameras for the acquisition of Earth imagery that can be downlinked to Earth. Cameras installed on the satellite include a visible spectrum camera and a near infrared imager. The satellite uses a VHF/UHF amateur radio terminal for telemetry downlink and command uplink while payload data uses an S-Band downlink system. All satellites subsystems have been designed from scratch and ANTELSAT will perform detailed testing of the various systems including the attitude control system that consists of a magnetometer unit and photodetectors for attitude sensing and magnetic torquers for coarse attitude control.

﻿Tigrisat﻿

Tigrisat is a 3U CubeSat developed at the University of Rome and launched aboard the UniSat-6 spacecraft for deployment after separation from the launch vehicle. The satellite carries an RGB imaging system using a dust detection algorithm to detect, monitor and study dust storm events over Iraq. A deployable VHF/UHF antenna system is used for command uplink and telemetry downlink while an S-Band system is used to downlink acquired data to two ground stations, one in Rome and one in Baghdad. Power is provided by body-mounted solar panels and nadir-pointing is accomplished through the use of magnetic coils and a b-Dot algorithm for attitude control.

﻿Duchifat-1﻿

Duchifat-1 – a 1U CubeSat – is an experimental and educational spacecraft developed at Herzliya Science Centre, Israel. The main objective of the mission is the transmission of real-time data from the satellite to ground stations using the amateur radio frequencies. The satellite will link to a number of LEO satellites and ground stations using the Automatic Position Reporting System (APRS) protocol. This will allow remote access for worldwide position/status reporting using simple mobile radios with omni-directional antennas. Real time position reporting is accomplished using GPS data. Duchifat-1 will be capable of transmitting a variety of data including short text messages and telemetry data acting as a digipeater. The APRS protocol is useful for SAR (Search and Rescue) and CAP (Civil Air Patrol) missions.

June 18, 2014

A Dnepr rocket is ready for liftoff from Site 370/13 of the Dombarovsky missile range at Yasny Launch Base, Russia, on Thursday at precisely 19:11:11 UTC to deliver a record-setting 37 satellites to orbit. This will break Dnepr’s previous record of 32 active satellites launched by a single rocket.

Over the past week, the payload integration process – which involves a high degree of complexity with a payload of 37 spacecraft – neared completion as all satellites were installed on their various adapters and deployers on the two payload platforms. Platform A hosts the Deimos-2, Tablesat-Aurora and BugSat-1 spacecraft and five QuadPacks containing 21 CubeSats contracted for launch through ISIS Launch Services. The second payload platform facilitates 13 satellites including KazEOSat-2, Saudisat-4, the Japanese Hodoyoshi-3 and 4, AprizeSat-9 and 10, BRITE-CA 1 & 2 and UniSat-6 which itself includes another four CubeSats. Detailed Overviews of all payloads can be found below.

When all payloads were installed, the two platforms were stacked and the Payload Fairing and Gas Dynamic Shield were attached. The integrated Space Head Module was moved to the Dnepr rocket silo on June 13 for installation atop the Dnepr rocket that had been inside the silo for several weeks undergoing checks and preparation. As part of the integration, the Space Head Module was structurally attached to the launch vehicle and data/electrical connections were made to allow communications between the controllers of the third stage and the payload deployment controller.

In the days leading up to launch, the first and second stage of the Dnepr were loaded with 149,900 Kilograms and 36,740kg of hypergolic propellants. Dnepr's third stage was fueled with 1,910kg of Unsymmetrical Dimethylhydrazine fuel and Nitrogen Tetroxide Oxidizer before installation on the launcher.

Final checkouts were on tap this week leading to the launch countdown that begins about eight hours before liftoff for the satellites and three hours before T-0 for the Dnepr launch vehicle. The silo-launched, three-stage Dnepr Launch Vehicle stands 34.4 meters tall and is 3.0 meters in diameter with a liftoff mass of 211,000 Kilograms.>>>Dnepr Launch Vehicle Overview

At T-0, the silo's mortar system will eject Dnepr that will immediately jettison its tray to start its first stage and soar into the sky headed for a high-inclination, near-circular Sun Synchronous Orbit at an altitude of 630 Kilometers and an inclination of 98.0 degrees.

Dnepr launches are usually not broadcast live. Confirmation of launch is normally published by Russian news agencies within minutes and confirmation of successful payload separation also arrives via those sources. Later on launch day, an official release with a launch photo or video is provided by launch operator Kosmotras.

After ejection from the Silo, Dnepr will ignite the RD-264 engine of the first stage which is a cluster of RD-263 engines sharing common turbopump equipment. With a total thrust of 461,200-Kilogram-force, Dnepr will start thundering uphill, starting its pitch-over to target the 98° insertion orbit.

Photo: ISIS Launch Services Weblog

Photo: ISIS Launch Services Weblog

Photo: Kosmotras

The first stage will burn until approximately T+1 minute and 38 seconds. Upon stage separation, the second stage ignites its RD-256 main and four-chamber RD-257 vernier engine to deliver a total thrust of 77,000-Kilogram force. During second stage flight, the upper portion of the payload fairing is jettisoned while the Gas Dynamic Shield still protects the payloads. The second stage will burn for about two minutes and 48 seconds before separating from the third stage.

*File Image* - Photo: Kosmotras

Before third stage ignition, the stack will coast to a point near the apogee of the trajectory so that the third stage burn can circularize the orbit. The third stage of the Dnepr launcher uses four RD-869 engines that are firing forward - towards the nose of the launch vehicle. This unusual propulsion scheme is a relic of the original R-36 ICBM concept that required the upper stage to perform multiple warhead deployments in a short time frame.

The engines deploy after staging and the vehicle makes a 180-degree flip to allow the third stage to provide a positive thrust to the stack, “pulling” the vehicle into orbit.

When the third stage, now firing in low-thrust mode, approaches the target orbit, the Gas Dynamic Shield will be jettisoned. Payload separation is completed in a carefully sequenced procedure starting with Deimos-2 and the rest of the Platform A payloads that are released as part of a pre-programmed sequence while the third stage is still firing & creating the proper separation distance between the payloads.

Once all payloads of Platform A are on their way, the platform itself will be jettisoned so that the 13 payloads of platform B can be deployed. The sequence of deploying all the satellites takes just about 30 seconds and all satellites end up in a slightly different orbit since the third stage continues firing for the separation sequence. After all payloads are deployed, the third stage continues firing to enter a graveyard orbit and avoid re-contact with the satellites.

﻿Dnepr Rocket to make record-setting Launch of 37 Satellites﻿

June 10, 2014

A Dnepr Rocket is being prepared to launch from the Dombarovsky missile range at Yasny Launch Base, Russia on June 19, 2014 carrying a total of 37 satellites to orbit to set a new record for most active satellites launched by a single launch vehicle - a record Dnepr itself is holding after the launch of 32 satellites of different shapes and sizes in November 2013.

Dnepr is a Russian/Ukrainian Launch System based on the R-36M Intercontinental Ballistic Missile that is now operated by launch provider ISC Kosmotras for orbital launches. The Dnepr Launch Vehicle stands 34.3 meters tall and is 3.0 meters in diameter with a liftoff mass of 211,000 Kilograms. The launcher has three stages in its basic configuration and can deliver payloads of up to 4,500 Kilograms to Low Earth Orbit.

When being inaugurated, Dnepr was designed for cluster missions – launching a large number of spacecraft of different sizes. For that, the vehicle uses different payload adapter systems and CubeSat dispensers.

Photo: ISIS Launch Services Weblog

Photo: ISIS Launch Services Weblog

Photo: GAUSS

Preparations for the mission began some time ago at the Yasny launch base when the Dnepr rocket was placed inside the silo at Site 370/13 for initial testing to ensure the launcher was functional. Also, the space head module was installed on the Dnepr rocket for customary fit checks before being moved to the integration facility where the 37 payloads were installed on the two decks of the space head module.

The KazEOSat-2 and Deimos-2 satellites as well as the larger secondary satellites were mounted on different payload adapters that fit the spacecraft and provide the necessary equipment for the separation which is controlled by a master timer and control unit that is also part of the space head unit. The various CubeSats were placed in different deployers including QuadPacks that support the timed launch of multiple small satellites. The payload integration campaign at the Yasny Launch Site is being completed this week as all 37 payloads find their place on the various payload adapters and dispensers for final checks to commence before the payload stack is installed atop the Dnepr launcher. In the days leading up to launch, the first and second stage of the Dnepr will be loaded with 149,900 Kilograms and 36,740kg of hypergolic propellants. Dnepr's third stage was fueled with 1,910kg of Unsymmetrical Dimethylhydrazine fuel and Nitrogen Tetroxide Oxidizer before installation on the launcher. Final checks will be completed next week, leading up to an eight-hour launch countdown for liftoff at 19:11:11 UTC on Thursday, June 19.

Photo: ISIS Launch Services Weblog

Photo: ISIS Launch Services Weblog

Photo: ISIS Launch Services Weblog

Photo: ISIS Launch Services Weblog

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